The present invention relates generally to coated gas turbine components, and more particularly components having airflow apertures and protective coatings.
Combustion chambers are engine sections which receive and combust fuel and high pressure gas. Gas turbine engines utilize at least one combustion chamber in the form of a main combustor which receives pressurized gas from a compressor, and expels gas through a turbine which extracts energy from the resulting gas flow. Some gas turbine engines utilize an additional combustion chamber in the form of an afterburner, a component which injects and combusts fuel downstream of the turbine to produce thrust. All combustion chambers, including both main-line combustors and afterburners, are constructed to withstand high temperatures and pressures.
Combustion chambers and other high-temperature gas turbine components vary greatly in geometry depending on location and application. All combustion chambers comprise a plurality of walls or tiles which guide and constrain gas flow, typically including a liner which surrounds a combustion zone within the combustion chamber. Liners and some other combustion chamber walls are conventionally ventilated with numerous air holes or apertures for cooling. Conventional apertures for this purpose are holes with walls normal to the surface of the liner. Some combustion chamber walls, including liners for main-line combustors and afterburners, receive thermal barrier coatings, coatings for erosion prevention, or radar absorbent coatings to reduce the radar profile of exposed portions of the turbine. Such coatings must withstand exceptionally high temperatures and pressures, and are frequently formed of brittle ceramics which are vulnerable to fracturing and delamination. Coatings in other high-temperature, high-pressure areas of gas turbines, particularly on combustor nozzles and hot turbine blades and vanes, share similar design requirements.
According to some prior art techniques, cooling apertures have been bored or punched in combustion chamber walls after coating deposition. More recent techniques apply coatings to combustion chamber walls and other gas turbine components after the formation of apertures. When using either technique, coatings near apertures are especially vulnerable to mechanical stresses, and are prone to fracture, ablate and delaminate from the substrate combustion chamber wall. A design solution is needed which reduces the stresses on combustion chamber wall coatings at aperture locations.